(1) Field of the Invention
The present invention relates to an aerofoil component of a gas turbine engine, and particularly an aerofoil portion which contains one or more passages for the transport of coolant therethrough.
(2) Description of Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
The performance of the simple gas turbine engine cycle, whether measured in terms of efficiency or specific output is improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbine at the highest possible temperature. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature will always produce more specific thrust (e.g. engine thrust per unit of air mass flow). However as turbine entry temperatures increase, the life of an un-cooled turbine falls, necessitating the development of better materials and the introduction of cooling mechanisms.
In modern engines, the high pressure (HP) turbine gas temperatures are now much hotter than the melting point of the blade materials used and in some engine designs the intermediate pressure (IP) and low pressure (LP) turbines are also cooled. During its passage through the turbine the mean temperature of the gas stream decreases as power is extracted. Therefore the need to cool the static and rotary parts of the engine structure decreases as the gas moves from the HP stage(s) through the IP and LP stages towards the exit nozzle.
Internal convection and external films are the prime methods of cooling the gas-path aerofoils, for example aerofoils, platforms, shrouds, shroud segments and turbine nozzle guide vanes (NGVs). Air is conventionally used as a coolant and is flowed in and around the gas-path aerofoils.
FIG. 1 shows an isometric view of a typical cooled stage of a gas turbine engine. Cooling air flows are indicated by arrows. FIG. 1 shows HP turbine NGVs 1 and HP rotor blades 2. Both the NGVs 1 and HP rotor blades 2 have aerofoil portions 100 which span the working gas annulus of the engine.
HP turbine NGVs generally consume the greatest amount of cooling air flow in high temperature engines. HP rotor blades typically use about half of the NGV cooling air flow. The IP and LP stages downstream of the HP turbine use progressively less cooling air flow.
The HP rotor blades 2 are cooled by using high pressure air from the compressor that has by-passed the combustor and is therefore relatively cool compared to the gas temperature. Typical cooling air temperatures are between 800 and 1000 K, while gas temperatures can be in excess of 2100 K.
The cooling air from the compressor that is used to cool the hot turbine components is not used fully to extract work from the turbine. Extracting coolant flow therefore has an adverse effect on the engine operating efficiency. It is thus important to use this cooling air as effectively as possible.
The ever increasing gas temperature levels combined with a drive towards higher Overall Pressure Ratios (OPR) and flatter combustion radial profiles, in the interests of reduced combustor emissions, have resulted in an increase in local gas temperatures and external heat transfer coefficients experienced by the HP turbine NGVs and rotor blades. This puts considerable demands on the internal and external cooling schemes that are heavily relied on to ensure aerofoil durability.
The last 10 years has seen a significant rise in the inlet gas temperature and overall engine pressure ratio on new engine designs, and this has brought a new raft of problems. However, the performance of the engine, and in particular the turbine, is still greatly affected by (a) the quantity of coolant consumed by the hot end aerofoils, and (b) the way the cooling flow is re-introduced into the gas-path. Therefore, while aerofoils must be provided with sufficient coolant flow to ensure adequate mission lives, it is imperative that the cooling scheme designs do not waste flow.
FIG. 2 shows a transverse cross-section through an HP turbine rotor blade aerofoil portion 100 with wall cooling around the suction surface S.
Suction side outer wall 110 and pressure side outer wall 140 define the external pressure side P and suction side S aerofoil surfaces of the aerofoil portion 100. Each outer wall 110, 140 extends from a leading edge LE to a trailing edge TE of the aerofoil portion 100. The aerofoil portion 100 in FIG. 2 has four main coolant passages 114 that extend in the annulus-spanning direction of the aerofoil portion 100. The front three of these passages are interconnected such that cooling air flows through the passages in series, reversing direction, as indicated by curved block arrows, between passages. The cooling air enters the main passages from feed passages at the root of blade, as indicated by the straight block arrows.
The aerofoil portion 100 further has a plurality of suction wall passages 106 that also extend in the annulus-spanning direction of the aerofoil portion 100. The suction wall passages 106 are bounded on opposing first sides by the suction side outer wall 110 and an inner wall 108 that separates the suction wall passages 106 from the main passages 114. Each suction wall passage 106 is bounded on opposing second sides by a pair of dividing walls 102 which extend between the suction side outer wall 110 and the inner wall 108. In each passage 106, one of the pair of dividing walls 102 is closer to the leading edge LE of the aerofoil portion 100 and the other of the pair of dividing walls 102 is closer to the trailing edge TE. Fillets 104 smooth the transitions from the dividing walls 102 to the inner wall 108 and to the suction side outer wall 110. As indicated by curved block arrows, coolant can flow in series through the suction wall passages with direction reversal.
However, this arrangement can cause thermo-mechanical structural problems and stress. A main cause of the stress results from differential thermal effects between the hot suction side outer wall 110 and the relatively cool inner wall 108, the highest thermal gradients occurring in the dividing walls 102 and fillets 104. For example, thermal growth of the hot suction side outer wall 110 is much greater than the cold inner wall 108 during transient throttle push, placing the outer wall 110 into compression and the inner wall 108 into tension. As a result, major stress concentrations are produced, particularly in the fillets 104. The thermal gradients at the dividing walls 102 further increase the overall stress levels. In particular, the fillet radii of the fillets 104 closest to the suction surface S are initially in compression during take off conditions when the suction side outer wall 110 reaches its maximum temperature. The local stress level in these fillets 104 can cause the material of the blade to plastically deform or creep such that when the suction side outer wall 110 cools down the fillets 104 can develop micro cracks in tension. When the process is repeated, cracks may propagate in the walls 102, 108, 110 due to low cycle thermal fatigue of the material.